1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a two piece turbine airfoil with cooling circuits.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially an industrial gas turbine engine, a compressor supplies compressed air to a combustor in which a fuel is burned to produce an extremely hot gas flow, which is then passed through a multiple stage turbine to extract mechanical energy used to drive an electric generator. The efficiency of the engine can be increased by passing a higher gas flow temperature into the turbine. However, the highest temperature of the turbine is limited to the material characteristics of the first stage stator vanes and rotor blades, since these are exposed to the highest temperature.
One method of allowing for a higher temperature on these first stage airfoils (blades and vanes) is to improve the cooling capabilities. Complex internal airfoil film cooling and impingement cooling circuits have been proposed to provide for improved cooling as well as to maximize the use of the cooling air.
FIG. 1 shows a prior art first stage turbine blade external pressure profile. The forward region of the pressure side surface experiences high hot gas static pressure while the entire suction side of the airfoil is at much lower hot gas static pressure than on the pressure side. FIG. 2 shows a prior art turbine blade with a 1+5+1 serpentine flow cooling circuit. The flow path for this 5-pass flow circuit is also shown in FIG. 1. The forward flowing 5-pass serpentine circuit is used in the airfoil mid-chord region. The cooling air flows toward and discharges into the high hot gas side pressure section of the pressure side. In order to satisfy the back flow margin criteria, a high cooling supply pressure is needed for this particular design, thus inducing high leakage flow. Since the second and third up-pass channels of the 5-pass serpentine cavities provide film cooling air for both sides of the airfoil, in order to satisfy the back flow margin criteria for the pressure side film row, the internal cavity pressure has to be approximately 10% higher than the pressure side hot gas side pressure which will result in over-pressuring the airfoil suction side film cooling holes. In the prior art 5-pass serpentine circuit, low aspect ratio flow channels are used. This lowers the ceramic core yield, making it difficult to install film cooling holes, high inference due to the rotational effect on internal heat transfer coefficient, and also yields low internal-to-hot gas side convective area ratio.
For the prior art highly cooled near wall turbine airfoil cooling design, the airfoil internal cavity has been used as cooling air passages or as a source cavity for supplying cooling air to various sections of the airfoil for the cooling system design. FIG. 3 shows a cross sectional view of a prior art near wall airfoil cooling design. A 5-pass counter flowing serpentine flow circuit is used in the airfoil mid-chord cavity to provide the airfoil tip section and suction wall cooling.
FIG. 4 shows another prior art near wall cooling design that utilizes the mid-chord cavity as a cooling source to provide cooling air to various sections of the airfoil for its cooling system design. Once the internal cavity is used as part of the cooling system, the inner wall of the double wall cooling structure will submerge in the coolant and become a cold structure membrane. Subsequently, it induces high thermal gradient for the dual wall cooling structure and yields a low component fatigue life. In addition, the airfoil internal cavity becomes a pressure vessel and a blade tip cap plus multiple internal ribs connecting the airfoil pressure side and suction side wall are required to ensure the structural integrity of the airfoil.
The above thermal mechanical fatigue (TMF) problem associated with the double wall turbine airfoil cooling design can be alleviated by incorporating the highly cooled turbine airfoil cooling design of the present invention into the prior art near wall airfoil cooling design.
It is an object of the present invention to provide for a turbine airfoil with a near wall serpentine flow cooling circuit which will optimize the use of main stream pressure gradient and two piece laminated blade construction.
It is another object of the present invention to provide for a turbine blade with near wall cooling that does not use a blade tip cap or internal cold ribs for support.